1. Field of the Invention
The present invention relates to gas turbine engines, and more specifically to cooling of turbine airfoils.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The hot gases flow downstream through turbine stages which extract energy therefrom for powering the compressor and producing useful work, such as powering a fan for propelling an aircraft in flight.
A high pressure turbine is disposed immediately downstream from the combustor and receives the hottest combustion gases from the combustor. The first stage turbine rotor blades have hollow airfoils which are supplied with a portion of air bled from the compressor for use as a coolant in removing heat from the blades during operation.
Each airfoil includes pressure and suction sidewalls joined together at opposite leading and trailing edges, and extending from root to tip. A platform is disposed at the airfoil root and defines a portion of the radially inner flow path for the combustion gases. And, a dovetail is integrally jointed to the platform for mounting the individual blades in corresponding dovetail slots in the perimeter of a rotor disk.
Since the airfoil leading edge first engages the hot combustion gases, it requires substantial cooling for obtaining a useful blade life. Heat load from the combustion gases varies around the outer surface of the airfoil from the leading to trailing edges, and along the pressure and suction sidewalls. Various cooling circuits are provided inside the airfoil for cooling the different portions thereof. The different portions of the airfoil therefore operate at different temperatures, which introduce thermal stress therein that affect low cycle fatigue life of the blade.
Airfoil cooling may be affected using convection cooling, film cooling, or impingement cooling, or combinations thereof. The leading edge of a first stage turbine airfoil typically includes several rows or columns of film cooling holes fed by a common leading edge flow chamber or channel. Other film cooling holes and trailing edge holes may be fed by corresponding internal channels, such as multi-pass serpentine cooling channels.
The airfoil may include additional film cooling holes disposed in either sidewall (pressure side or suction side) downstream of the leading edge, which are typically referred to as gill holes. Since the gill holes are typically provided with a common source of coolant inside the airfoil, and the pressure of the combustion gases outside of the airfoil varies, backflow margin across the gill holes may vary on opposite sides of the airfoil.
Backflow margin is defined as the pressure of the coolant inside the airfoil divided by the local pressure of the combustion gases outside the airfoil as experienced by each of the gill holes. Sufficient backflow margin must be maintained to prevent ingestion of the hot combustion gases into the airfoil, and ensure continuous discharge of the coolant through the gill holes.
Since the minimum required backflow margin must be set at the airfoil leading edge pr pressure sidewall, the backflow margin on the lower suction sidewall of the airfoil may be undesirably high.
FIG. 1a shows a typical Prior Art (1+3) serpentine cooling design for the first blade of the turbine. The flow path for the 3-pass flow circuit is also shown in FIG. 1b. the airfoil includes a first leading edge cooling passage 101, film cooling holes 102 to deliver cooling air from the leading edge cooling passage 101 to a second leading edge cooling passage 103, a 3-pass serpentine passage having a first leg 104, a second leg 105, and a third leg 106, and trailing edge film cooling passages 107 supplied by cooling air from the first leg 104 of the serpentine passage. For a forward flowing 3-pass serpentine cooling design used in the airfoil mid-chord region, the cooling air flows toward and discharges into the high pressure hot gas side pressure section of the pressure side of the blade. In order to satisfy the back flow margin criteria, a high cooling supply pressure is needed in order to prevent the hot gasses from flowing into the airfoil.
Since the last leg of the 3-pass serpentine cavities provides film cooling air for both sides of the airfoil, in order to satisfy the back flow margin criteria for the pressure side film row, the internal cavity pressure must be approximately 10% higher than the hot gas pressure of the pressure side of the airfoil. The high pressure required preventing inflow from the high pressure side of the airfoil (the pressure side) results in an over-pressuring of the airfoil suction side film holes since the film cooling holes of the pressure side and the suction side is connected to the same cavity.
The U.S. Pat. No. 6,168,381 B1 issued to Reddy on Jan. 2, 2001 and entitled AIRFOIL ISOLATED LEADING EDGE COOLING discloses a serpentine cooling passage design in which an isolation flow chamber (38 in FIG. 3 of this patent) is positioned between a pressure side and suction side flow channels (40 and 42 in FIG. 3), where the pressure side and suction side flow channels are the last leg in a 3-pass serpentine flow circuit, both being supplied with cooling air from a common first and second legs of the 3-pass serpentine circuit. Because both pressure side and suction side flow channels are supplied from the same upstream cooling air passage, the pressures in the pressure and suction side flow channels are the same. The same problem described above exists in the Reddy patent: a high pressure is required to prevent inflow of the hot gasses on the pressure side of the airfoil, and the suction side channel is over-pressurized resulting in excessive flow through the film cooling holes on the suction side of the airfoil.
U.S. Pat. No. 5,813,835 issued to Corsmeier et al on Sep. 29, 1998 and entitled AIR-COOLED TURBINE BLADE is another Prior Art design to improve airfoil cooling by separating the pressure side serpentine passage from the suction side serpentine passage. The Corsmeier patent (reproduced in FIGS. 2a and 2b of this disclosure) has a pressure side 3-pass serpentine cooling circuit 222 that flows from the trailing edge side of the airfoil toward the leading edge side, and a suction side 3-pass serpentine cooling circuit 224 that also flows from the trailing edge side toward the leading edge side. The Corsmeier patent also includes a leading edge cooling passage 228 (28 in FIGS. 4-8 of this patent) and a trailing edge cooling passage 230 (30 in FIGS. 4-8) for a total of four separate cooling passages in this patent. One problem with the Corsmeier patent is that the leading edge cooling passage 228 and the trailing edge cooling passage 230 are not continuous passages, but are dead-end passages which results in poor cooling air flow toward the end or tip of the passage. Most of the cooling air flows out of the passage before reaching the tip. Therefore, inadequate cooling results near the blade tip.